A gas turbine engine includes one or more turbine blade rows disposed downstream of a combustor which extract energy from combustion gases generated thereby. Disposed radially outwardly of the rotor blade tips is a stator shroud which is spaced from the blade tips to provide a relatively small clearance therebetween for reducing leakage of the combustion gases over the blade tips during operation. Each of the rotor blades includes conventionally known pressure and suction sides which are preferentially aerodynamically contoured for extracting as much energy as possible from the combustion gases flowable thereover. The pressure and suction sides extend to the blade tip and are disposed as close as possible to the stator shroud for maximizing the amount of energy extracted from the combustion gases. However, the clearance between the blade tips and the stator shroud must nevertheless be adequate to minimize the occurrence of blade tip rubs during operation which may damage the blade tips.
Turbine rotor blades are typically hollow for channeling therethrough cooling air which is provided from a conventional compressor of the gas turbine engine to cool the blades from the heat flux generated by the combustion gases flowing thereover. The tip, or tip cap, portion of the blades is particularly susceptible to the damaging effects of the hot combustion gases and must be suitably cooled for reducing blade tip distress in the form of oxidation and thermal fatigue during operation. As the blade tip erodes during operation due to the blade tip distress, the pressure and/or suction sides of the blade are adversely affected which decreases the aerodynamic performance efficiency of the blade for extracting energy from the combustion gases. And, such erosion of the blade tip also increases the clearance between the blade tip and the stator shroud which allows more of the combustion gases to leak over the blade tip, and, therefore, extraction of the energy therefrom is lost which also decreases aerodynamic efficiency.
Numerous conventional blade tip cap designs exist for maintaining the proper pressure and suction side flow surfaces of the blade at the tip cap as well as providing minimum clearances with the stator shroud. Numerous cooling configurations also exist for cooling the blade tip caps for obtaining useful lives of the blades without undesirable erosion thereof. Since cooling of the blade, including the blade tip, uses a portion of the compressed air from the gas turbine compressor, that air is unavailable for combustion in the combustor of the engine which decreases the overall efficiency of the gas turbine engine. Accordingly, cooling of the blade, including the blade tip, should be accomplished with as little compressed air as possible to minimize the loss in gas turbine engine efficiency.
Conventional blade cooling configurations typically include film cooling apertures extending through the blade pressure and suction sides for channeling the cooling air from inside the blade to the outer surfaces of the blade for providing conventional film cooling thereof. The film holes must be suitably aligned with each other on the blade outer surface which is typically accomplished by using a conventional reference datum located on the blade dovetail. However, manufacturing tolerances and stack-up tolerances result in random inaccuracy in position of the film cooling holes adjacent to each other which affect the ability to provide the designed-for cooling thereof. Yet further, the film cooling holes typically extend at inclination angles through the blade walls which result in elliptical outlet holes along the outer surface of the blade. In conventional radially inclined holes, the major axis of the film cooling outlet also extends radially which provides an undesirable stress concentration of the principle thermal stresses within the blade which extend generally in the axial direction, for example. These stress concentrations must be suitably accommodated in the design of the rotor blade for obtaining a useful service life thereof.